(1) Field of the Invention
The present invention relates to a rotary wing aircraft having a non-ducted tail rotor with at least five blades. The invention thus relates to the technical field of rotorcraft tail rotors.
(2) Description of Related Art
A rotorcraft is usually provided with at least one main rotor providing the rotorcraft with at least part of its lift, and possibly also with its propulsion. In particular, and by way of example, a helicopter may include a single rotor providing it both with lift and with propulsion. Because of this function, such a rotor is referred to as a “main” rotor.
The main rotor may be driven by a power transmission gearbox, itself driven by at least one engine. In its rotation, the main rotor induces torque on the fuselage of the helicopter. This torque then tends to cause the fuselage to perform turning movement about a yaw axis. Consequently, a helicopter that has a single main rotor is usually provided with a device for controlling its yaw movement.
Such a device may comprise an auxiliary rotor arranged at the tail end of the helicopter. The auxiliary rotor exerts thrust having at least a transverse component for the purpose of controlling the yaw movement of the aircraft. Under such circumstances, the auxiliary rotor is sometimes referred to as the “tail” rotor because of where it is arranged, or indeed as an “anti-torque” rotor because of its ability to counter the torque generated by the main rotor on the fuselage. The term “tail” rotor is usually used below.
In order to control the yaw movement of the aircraft, a pilot controls the collective pitch of the blades of the tail rotor. Modifying this collective pitch leads to a modification to the thrust generated by the tail rotor. For example, a pilot may operate pedals connected to the tail rotor in order to control the movement of the aircraft about its yaw axis.
In a first alternative, the tail rotor is a non-ducted rotor.
Such a non-ducted rotor has a hub that is transversely offset relative to a tail fin. The hub carries at least two blades of large span. The blades present profiles with chords that are relatively large.
Thus, a light helicopter sometimes has a hinged tail rotor with two blades. A heavy helicopter may have at least five blades. The number of blades is determined as a function of the thrust that needs to be delivered in order to control the yaw movement of the aircraft. A manufacturer tends to minimize this number of blades in order to optimize the number of parts of the rotor, their price, and their weight.
Document U.S. Pat. No. 2,818,224 describes an aircraft having a non-ducted tail rotor carried by a vertical fin.
Furthermore, the Ecureuil® helicopter has two blades connected together diametrically by a common spar. The spar is hinged to a drive shaft. Thus, two half-shells surround the spar. The half-shells are hinged to a clevis of the shaft by a pin. Such a tail rotor is referred to as a “hinged” rotor because it is hinged to a drive shaft.
The spar has two twistable zones arranged on either side of the half-shells. Consequently, each twistable zone can be twisted to allow the pitch angle of each blade to be modified.
In order to control the pitch angle of a tail rotor blade, the aircraft has one pitch rod per blade. Each pitch rod is hinged to the root of the blade. For example, a collar is secured to the cover of the root of a blade, the collar then being hinged to a pitch rod.
Furthermore, laminated bearings are arranged between the half-shells and the collars. Such bearings are deformable in twisting and in shear.
Flight controls are consequently connected to the pitch rods in order to control the turning of each blade about a pitch axis. Such turning twists the twistable zones of the spar.
In addition, each blade is subjected in flight to aerodynamic forces.
The blades of a non-ducted tail rotor move through an air stream that is generated by the helicopter advancing. This air stream is referred to below as the “longitudinal” air stream, with reference to the direction of longitudinal advance of the aircraft.
Consequently, in its rotation, each blade travels over a half-revolution in the advancing direction of the helicopter in a position referred to as the “advancing” blade position. Conversely, each blade rotates through another half-revolution in the direction opposite to the advance direction of the helicopter in a position referred to as the “retreating” blade position.
Consequently, an advancing blade travels at a speed equal to the sum of the speed of rotation of the tail rotor plus the forward speed of the aircraft. Conversely, a retreating blade travels at a speed equal to the difference between the speed of rotation of the tail rotor and the forward speed of the aircraft. Unfortunately, the lift generated by a blade varies as a function of its speed relative to the incident air stream. Under such circumstances, a blade generates smaller lift in its retreating blade position than in its advancing blade position.
Since the tail rotor is hinged, this difference in lift leads to the tail rotor tilting about a substantially vertical tilt axis. This tilt axis corresponds to the axis of the pin hinging the half-shells of the rotor to a drive shaft.
This dynamic tilting of the tail rotor imparts dynamic forces to the blade pitch control linkage.
In particular, the pitch rod of a blade may be radially offset relative to a flapping axis of the blade, i.e. the axis of the pin in the above-described tail rotor context.
Consequently, the tilting of the rotor then tends to exert a dynamic force on each pitch rod.
In addition, static forces are also imparted to the blade pitch control linkage.
Each blade tends to be subjected to a return moment caused by the centrifugal force exerted on the blade. This return moment has the effect of returning each blade towards the plane of rotation of the rotor. This return-to-flat moment leads to forces being created in each pitch rod.
Consequently, the rotation of a blade leads to static and dynamic forces being created in the blade pitch control linkage. A pilot can then have difficulty in modifying the pitch of the blades, e.g. by moving pedals.
In order to remedy that, a servo-control may be interposed in the control linkage. The pilot's flight controls are then connected to the servo-control and it is the servo-control that is connected to each pitch rod.
Nevertheless, the provision of a servo-control can lead to a non-negligible increase in weight.
Furthermore, in the event of a hydraulic failure, the servo-control behaves like a rod. In order to assist the pilot in this configuration, the aircraft may include a force compensator. A force compensator may include a hydraulic pressure accumulator and a control lever interposed between the pitch rods and the servo-control.
In order to reduce the forces generated on a pitch control linkage, and in particular the static forces, the root of each blade may be provided with two heavy members referred to as “Chinese” weights.
Each Chinese weight projects in elevation from the cover of a blade in order to create a moment that tends to oppose the return moment. Nevertheless, Chinese weights have limited effectiveness on the dynamic forces suffered in flight, and can even tend to increase such dynamic forces.
Under such circumstances, a non-ducted tail rotor may be a hinged rotor that is subjected to static forces and dynamic forces that can stiffen the flight controls. Furthermore, these forces can be penalizing on the lifetime of laminated bearings, for example.
In addition, a non-ducted tail rotor is usually carried by a tail fin of small dimensions. Consequently, in the event of a break in the power transmission driving the tail rotor in rotation, the tail fin runs the risk of generating insufficient lateral lift to counter the torque generated by the main rotor.
In a second alternative, the tail rotor is a ducted tail rotor.
A ducted tail rotor includes a fairing. The fairing is a component portion of a tail fin of large dimensions.
In addition, a ducted tail rotor has a rotor of small dimensions arranged within the fairing. The rotor is provided with a hub carrying small-span blades of profiles that present chords that are relatively small.
The thrust generated by the rotor is then relatively small, about half the thrust generated by a non-ducted tail rotor. Nevertheless, the fairing also generates thrust that is about half of the thrust generated by a non-ducted tail rotor. Under such circumstances, the fairing and the rotor of a ducted tail rotor together generate total thrust that is sufficient for controlling the yaw movement of the aircraft.
Installing ducted tail rotor is advantageous, but nevertheless requires a fin of large dimensions to be put into place. Such a fairing is then penalizing from a weight point of view.
In addition, the fairing protects the rotary portion of a ducted tail rotor against the impact of a longitudinal air stream. Consequently, a ducted tail rotor avoids creating dynamic forces by arranging a fairing around the blades.
Document FR 2 719 554 describes a ducted tail rotor.
Document FR 2 699 497 describes a device for connecting a blade to a hub that is applicable to the blades of a variable pitch multiblade rotor, for a ducted tail rotor of a helicopter.
Documents that do not belong to the technical field of the invention are mentioned by way of information only.
Document US 2012/0219417 describes a rotor with retractable blades.
Document US 2004/0113013 describes a lift rotor having a peripheral ring.
Document U.S. Pat. No. 4,913,376 describes two contrarotating rotors of an autogyro.
Document U.S. Pat. No. 4,195,800 describes a lift rotor for an autogyro, the rotor having a central disk of large dimensions.
Documents FR 1 411 762 and EP 0 018 114 are also known.